The present invention relates generally to rocket propulsion systems and specifically to hybrid rocket engines. There are three basic types of chemical rockets in use today: liquid rocket engines that use liquid propellants, solid rocket motors that use solid propellants, and hybrid rocket engines that use a combination of liquid and solid propellants.
In a conventionally designed hybrid rocket engine, the fuel is stored in the solid state, while the oxidizer is stored in either the liquid or gaseous state. Traditionally in most hybrid rocket engine designs, the solid fuel is cast-molded, extruded, or in some instances machined into a cylindrically shaped structure referred to as a fuel grain. The fuel grain is designed and formed to feature one or more internal passages running through its length. These passages are referred to as ports. The fuel grain port or ports dually serve as the hybrid rocket engine's combustion chamber or chambers, and through a gas phase change and ablation process, the fuel source.
The fuel grain is conventionally housed within a metal or fiber-reinforced polymer composite motor case designed to withstand the pressures and elevated temperatures created during the combustion process. The motor case may also feature an internal liner made from a high-temperature material to create a thermal barrier to prevent damage or burn-through during the rocket engine's operation.
The motor case, with fuel grain installed, is attached to a forward cap typically machined or cast from high-temperature metal alloys. The forward cap forms the pre-combustion chamber and houses the oxidizer injectors and ignition system. The aft end of the motor case is attached to an assembly which forms the post combustion chamber and allows secure attachment to the rocket nozzle. The assembled motor case with fuel grain installed, forward cap, and aft assembly with attached nozzle is conventionally referred to as the motor or solid section of the hybrid rocket engine.
In a conventionally designed hybrid rocket engine, liquid or gaseous oxidizer is stored separately in an integrally formed pressure vessel or tank forward of the motor section within the rocket powered vehicle. However, in some designs, liquid or gaseous oxidizer may be stored adjacent to the motor section or even remotely on the vehicle. Conventionally, the tank or pressure vessel stored liquid or gaseous oxidizer is urged through a specially designed plumbing system, typically including a flow control valve to feed oxidizer through one or more oxidizer injectors housed within the motor section forward cap; and in turn, through the fuel grain port or ports.
The motive force needed to urge the liquid or gaseous oxidizer through the oxidizer injector or injectors into the fuel grain port or ports with sufficient flow rate to support combustion may be generated by any one of several means such as enabling a liquid to gas phase change, causing an exothermic reaction using a catalyst, employing a mechanical boost pump, pre-pressurizing the oxidizer tank with an externally supplied inert gas, or using an on-vehicle high pressure tank filled with an inert gas to boost oxidizer tank pressure.
Regardless of the configuration or type of liquid or gaseous oxidizer used, the assembly of oxidizer tank, pressurizing system and associated plumbing is typically referred to as the oxidizer section. Collectively, the motor section and the oxidizer section are referred to as the hybrid rocket engine, sometimes also referred to as the hybrid rocket motor.
Hybrid rocket engines offer certain advantages over both solid rocket motors and liquid rocket engines alike. For example, once ignited, a solid rocket motor cannot be stopped until its propellant is exhausted and it cannot be throttled or restarted. Hybrid rocket engines, like liquid rocket engines, can be designed for on-command thrust termination, throttling, and engine restart. Most liquid monopropellant rocket engines use highly toxic, environmentally damaging propellants that are now considered too dangerous and to environmentally unsafe for continued use.
Compared to most liquid bi-propellant rocket engines, hybrid rocket engines are significantly less mechanically complex, and therefore more reliable and less expensive to develop, manufacture, and operate. Hybrid rockets are ideally suited to use propellants that are self-pressurizing, non-toxic, environmentally benign, operate at ambient temperatures, and require no specialized equipment for handling, transporting, and loading. Furthermore, hybrid rocket engines, due to their propellants being stored in different states of matter, are inherently immune to explosion. Immunity to explosion is of great importance to rocket-powered vehicle designers and operators. Their superior safety, mechanical simplicity compared to liquid bi-propellant rocket engines, and environmental friendliness all translate to improved reliability as well as lower development, manufacturing, and operating costs.
Despite all of their aforementioned advantages, conventionally designed hybrid rocket engines using cast-molded solid fuels like hydroxyl-terminated polybutadiene (HTPB), a form of synthetic rubber that has been the most studied hybrid rocket engine fuel to date, are rarely if ever employed for applications requiring vibration free, consistent high performance. Unfortunately, conventionally designed hybrid rocket engines using cast-molded HTPB as well as other cast-molded solid fuels, including paraffin wax, polyamides, and thermoplastics have not been able to demonstrate the vibration free, consistent, high performance required for most rocket propulsion applications.
Excessive vibration and inconsistent performance is even more pronounced when higher energetic additives such as aluminum powder have been blended into solid fuels like HTPB and paraffin wax. All of these disadvantages and inefficiencies are attributable to either the solid fuel material selected or the fuel grain production methods used. To fully understand the efficacy and advantages of the present invention, it is important to understand these disadvantages in relation to competing rocket propulsion systems as well as their respective causes.
Comparative poor hybrid rocket engine performance and their often unpredictable, even sometimes dangerous nature can be attributed to: 1) low regression rate, i.e., the rate at which the solid fuel is consumed compared to solid rocket motors, 2) adverse harmonics build-up inducing unacceptable, sometimes dangerous levels of vibration, 3) excessive solid fuel waste compared to other rocket propulsion systems, 4) low specific impulse (Isp) compared to most liquid bi-propellant rocket engines, and 5) inconsistent, unpredictable thrust performance which renders them unusable in clustered (multiple engines per launch vehicle stage or spacecraft) configurations.
1). Low Regression Rate. For a given selection of fuels and oxidizer-to-fuel mass ratios, the thrust generated by a rocket or any type of reaction engine is approximately proportional to the mass flow rate. In a hybrid rocket engine, mass flow rate is proportional to fuel grain regression rate. In a classically designed hybrid rocket engine, particularly those using slow burning fuels like HTPB, the burning rate is further limited by the heat transfer from the relatively remote flame to the fuel grain port surface. One of the physical phenomena that limit the burning rate is the blocking effect that is caused by the injection of vaporizing fuel into the high-velocity oxidizer gas stream. Given the linear nature of the oxidizer gas stream, oxidizer/fuel vapor mixing and resulting combustion efficiency is a function of the amount of time available for mixing to occur within a classically designed hybrid rocket fuel grain port.
Attempts to increase the burning rate by mixing energetic materials like Alcoa produced Military Grade 44 aluminum powder (Rockledge, Tex.) (average particle size of 44 microns) with traditional hybrid rocket fuels using cast-molding production methods have been only marginally successful in improving rocket engine performance. Aluminum powder is highly reactive with oxygen and water. To passivate the material to become stable in atmospheric conditions for safe handling, processing, storing, transporting, and use in a rocket engine, the aluminum particle is allowed to form an outer layer of aluminum oxide (alumina), a non-combustible material that when burned acts as a heat sink causing a loss of temperature and energy within the center port.
Nano-scale Aluminum powder is thought to be the next big advancement in both solid and hybrid rocketry. Elemental Aluminum in nano-scale is significantly higher in reactivity than micron-scale powder due to its relatively high specific surface area. Unfortunately, most attempts to safely and efficaciously employ this material in both solid and hybrid rocketry have not been successful. If allowed to form an alumina shell, effectively consuming a portion of the aluminum core, much of the elemental aluminum's energetic value is lost.
In addition to the challenges associated with obtaining a uniform blend of polymer and metal powder throughout the fuel grain using the cast-molding technique, improved burning rates by use of metal additives such as aluminum have only served to exacerbate the problems associated with using relatively elastic materials such as HTPB and paraffin waxes as a primary hybrid rocket solid fuel. Moreover, attempts to improve on regression rate further using high energetic material such as ALEX powder (an ultra-fine aluminum powder produced by the plasma-explosion process) have been even less successful and have introduced a significant potential for spontaneous ignition or explosion stemming from the pyrophoric nature of these ultra-fine powders.
Despite the potential for significant increase in burning rate, on the order of 30% higher than standard Military grade 44 micron particle size aluminum powder, employing a material that will spontaneously ignite upon exposure to the atmosphere or explode on contact with water or water vapor is counter-productive to one of the most significant advantages of a hybrid rocket engine—its comparative higher safety (i.e., benign failure mode and U.S. Government recognized zero TNT equivalency) compared to other forms of chemical rocketry.
More recent efforts have involved the development of methods to stabilize the nano-scale aluminum particles by encapsulating each particle in a polymeric material; thereby, protecting the elemental aluminum from the environment. While some of these approaches such as emersion in benzene followed by compounding with styrene to form granules of aluminum-styrene have merit and warrant further investigation, the breakthrough developed by St. Louis, Mo. based NanoMetallix, LLC is of note and interest. This firm has developed a process in which the elemental aluminum particle, measuring 15 nm or less is produced in a reactor simultaneously with the formation of a crystalline polymer outer shell. Thus, the NanoMetallix passivated material is not only safe to handle, transport, store, and use as rocket propellant, the particle core remains 99.9% pure elemental Aluminum.
This difference in the combustion scheme of a hybrid rocket engine significantly degrades the propellant burning rate compared to a solid rocket motor propellant in which the solid state oxidizer and fuel are in intimate contact. Consequently, the regression rate, using conventionally molded fuel grain materials like HTPB is typically one-tenth or less than that of most solid rocket propellants.
Structurally soft, HTPB with a Young's Modulus varying between 0.0026 GPa and 0.00756 GPa is a common polymeric binder used in solid rocketry. It has been the fuel of choice for over fifty years in many U.S. Government sponsored hybrid rocket propulsion research projects. Most of this work has involved integrating multi-port configurations into the fuel grain's design to increase the total fuel grain port surface area as a means to improve regression rate. Unfortunately, improvements in regression rate using multi-port designs have been offset by reduced fuel volume loading, adverse harmonics built-up that induces excessive and sometimes dangerous levels of vibration, unpredictable thrust performance, and increased fuel waste. However, excessive vibration, unpredictable thrust performance, and increased fuel waste have also been observed in single port large hybrid rocket engine designs using both HTPB as well as faster burning, also structurally soft, paraffin wax with a Young's Modulus of 0.061 GPa. While it is generally understood that regression rate in a hybrid rocket engine is a function of fuel burn rate and port surface area, the increased regression rates achieved using multi-port grain configurations have been more than offset by reduced reliability, consistency, efficiency, and safety.
2) Adverse Harmonics and Excessive Vibration. In any discussion about vibration in a hybrid rocket engine, it is important to keep in mind that the port within a hybrid rocket fuel grain is the engine's combustion chamber. Combustion chamber wall integrity is an essential design criterion in any reaction engine. Therefore, it is understandable that if a combustion chamber wall's structural integrity is degraded or compromised, chamber performance and reliability would likewise be degraded or compromised. Logically, an engineer would be reluctant to use a compressible, easily fractured material to fabricate a combustion chamber. But, this is exactly the case when soft, compressible, and fracture prone materials like HTPB and paraffin wax are used to construct a hybrid rocket fuel grain and its combustion chamber port or ports. To make matters more complex, given the fuel grain is also the rocket engine's fuel supply, as fuel is consumed, the port wall continually ablates and expands in diameter; thereby, increasing available surface area causing an oxidizer-fuel mixture shift from oxidizer rich to fuel rich combination. Materials such as HTPB and paraffin wax are thought to respond to high pressure gases created within the port by compressing the solid fuel against the higher-strength motor case; thereby, inducing grain fractures and erosive burning—both common occurrences in large scale HTPB and paraffin wax hybrid rocket engines.
Adverse harmonics exhibited in hybrid rocket engines, particularly pronounced in large-scale variants, is thought to be caused by a compressive-relaxation response by these soft fuels reacting to elevated chamber pressures, creating a type of trampoline effect. These oscillations can build to dangerous vibration levels and even a catastrophic over pressurization event. Cast-molded fuel grains made from these materials are also prone to structural flaws such as weak spots, air bubbles, hot spots, and fractures that are also known to cause erosive burning and erratic, unpredictable performance. Fuel fragments breaking free and blocking or temporarily blocking the rocket's nozzle have also been recorded. These phenomena are considered even more problematic in large hybrid rocket engines, especially those using multi-port designs.
3). Excessive Solid Fuel Waste. A certain amount of residual solid fuel is expected in a hybrid rocket engine. However, in a multi-port configuration, the amount of non-combusted fuel that is expelled can be significant and in certain circumstances a safety concern. In multi-port designs, as the burn progresses and fuel is ablated and combusted, the structure between the ports ultimately losses its integrity until failure occurs. In these situations, chunks of non-combusted fuel and webbing material have been known to break free, partially and sometimes completely blocking the nozzle, which can cause a serious safety problem. In multi-port HTPB fueled hybrid rocket engine designs, the total amount of residual and unspent fuel can reach 15% or more.
4). Poor Specific Impulse. Expressed in seconds, specific impulse (usually abbreviated Isp) is a measure of the efficiency of rocket and jet engines. By definition, it is the total impulse (or change in momentum) delivered per unit of propellant consumed and is dimensionally equivalent to the generated thrust divided by the propellant flow rate. Typically referenced as performance in vacuum for rockets, Isp is a convenient metric for comparing the efficiency of different rocket engines for launch vehicles and spacecraft.
Generally speaking, there is an inverse relationship between increased regression rate and Isp in a hybrid rocket. Whereas, regression rate speaks to the hybrid rocket engine's volumetric efficiency and thrust output as a function of fuel grain diameter, Isp relates more to the rocket engine's propellant efficiency. Ideally, rocket engine designers attempt to improve both. However, attempts to improve on hybrid rocket Isp has mainly focused on evaluating and testing different propellant combination. Whereas, a classical hybrid rocket engine uses a liquid or gaseous oxidizer and solid fuel, past experiments have been conducted on engine's that use a solid oxidizer and liquid fuels. While many of these achieved very high Isp—in the high 300 seconds (vacuum), they proved to be impractical for reasons mostly associated with the need to maintain a hydrocarbon fuel as a solid at cryogenic temperatures.
Other approaches have involved blending energetic materials such as aluminum powder into the fuel grain composition to increase Isp. However, obtaining a consistent, uniform mixture has always been a challenge using cast-molding techniques, especially when molding multi-port grains. Most conventionally designed hybrid rocket engines using nitrous oxide and polymeric fuel like HTPB average Isp is between 270 seconds to 290 seconds (vacuum), the higher figure attained with the addition of aluminum powder as an additive. While higher than most solid rocket motors, this level of performance is significantly lower than competing liquid bi-propellant systems using liquid oxygen and hydrocarbon fuels like kerosene that average between 310-340 seconds.
5). Inconsistent Thrust Performance. Inconsistent, unpredictable thrust in a classical hybrid rocket engine is a direct consequence of all of the above listed shortcomings and problems. Inconsistent and unpredictable performance makes it impossible for a hybrid rocket engine to be seriously considered for most rocket propulsion applications and uses. Further, many of the causes of inconsistent thrust performance can be tied to the cast-molding production process used to fabricate hybrid rocket fuel grains. HTPB and paraffin wax fuel grains are typically centrifugally cast-molded, with the latter containing a small percentage of polyethylene to improve tensile strength. During the HTPB polymerizing process, small air bubbles are formed and hot spots are created due to incomplete mixing and uneven curing. HTPB fuel grains require up to 90 days or more to fully cure, and even then, their material characteristics change over time. Small air bubbles are also formed during the cooling cycle when fuel grains are cast from paraffin wax. Bubble formation is a function of the shrinkage occurring within the wax. In an attempt to reduce or eliminate unwanted air bubbles as well as other types of grain flaws and hot spots, centrifugal casting methods, taking up to 120 hours to complete, are routinely employed. Even with these measures, air bubbles, structural cracks, hot spots, and other flaws seem to be chronic for fuel grains made using the cast-molding process.
Therefore, it would be highly desirable to develop a solid fuel propellant and fuel grain architecture-topology that exhibits: 1) flawless composition, 2) a regression rate comparable to solid rocket motors, 3) significantly improved thrust consistency, 4) more thorough oxidizer-fuel mixing, 5) greatly improved specific impulse, and 6) minimal vibration—all without compromising the many safety, mechanical simplicity, and economic advantages inherent in hybrid rocket propulsion systems.